This invention relates in general to fixed wing aircraft and in particular to fixed wing aircraft in which the aircraft is provided with structural elements that allow the wing to be rotated about a generally vertical axis. In the prior art devices, the wing pivot structure is usually intended to allow rotation of the wing during flight in order to change the flight angle of the leading edge of the wing with respect to the line of flight of the aircraft. In these applications of wing rotation devices, the angle of rotation is limited to a few degrees. Examples of these structures are shown in U.S. Pat. Nos. 3,510,088, 3,971,088 and 4,132,374.
The present invention is intended for use in an aircraft in which the wing is to be rotated only in non-flight conditions. In particular, the invention is described in connection with its use on a fixed wing aircraft intended for use on an aircraft carrier where it is necessary to change the wing position in order to reduce the overall size of the aircraft to allow it to be carried on an elevator to the lower hangar decks of the ship and to minimize the hangar space required for storage of the aircraft.
In addition to providing means for rotation of the wing, the rotating structure must also carry wing flight loads and, more importantly, it is most desirable that the structure also absorb the wing box strains resulting from flight loads rather than require the fuselage airframe to absorb these reversing motions. In order to absorb wing box strains, conventional wing pivot structures of the prior art require a plurality of links between the wing box and the fuselage which are provided with spherical or universal bearings at their ends to allow the wing box to flex relative to the fuselage. This series of linkages adds considerable weight to the aircraft and increases its complexity. Increased complexity introduces multiple wear points, thus generally reducing overall the reliability of the entire aircraft.
The present invention overcomes these disadvantages by attaching the wing to the fuselage by means of a simple unitary ring which is assembled to form an integrated structure with the wing box structure. The ring is allowed to flex at its wing attachment points an amount sufficient to absorb the expected wing box strains. The ring is attached to the fuselage structure by four primary attachments assemblies which may be disconnected from the ring to allow wing rotation. The majority of the vertical flight loads are carried directly by the column-like supports which are integrated into the unitary ring between the primary wing and fuselage attachment points. Secondary attachment assemblies on the fuselage provide additional in-plane flight load paths and support for rotation of the ring by means of shoes attached to the fuselage and connected in sliding engagement to the ring. In-plane loads are reacted by the ring in a tangential direction in which the ring is inherently very stiff as compared to the radial flexibility of the wing attachment points of the ring.
Thus it can be seen in the following detailed description of the invention, that there is herein described a structure for allowing rotation of an aircraft wing during non-flight conditions which requires a minimum of parts thus reducing the aircraft overall weight while increasing the reliability of the entire machine.